The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and the introduction of internal air cooling.
In modern engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these aerofoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.
Internal convection and external fluid films are the prime methods of cooling the gas path components—aerofoils, platforms, shrouds and shroud segments etc. Cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.
Ever increasing gas temperature levels combined with a drive towards flatter combustion radial profiles, in the interests of reduced combustor emissions, have resulted in an increase in local gas temperature experienced by the extremities of the blades and vanes, and the working gas annulus end-walls.
A turbine blade or vane has a radially extending aerofoil portion with facing suction side and pressure side walls. These aerofoil portions extend across the working gas annulus. Cooling passages within the aerofoil portions of blades or vanes are typically fed cooling air by inlets at the ends of the aerofoil portions. Cooling air eventually leaves the aerofoil portions through exit holes typically positioned at the trailing edges and, in the case of blades, the tips. Some of the cooling air, however, can leave through film cooling holes formed in the suction side and pressure side walls.
FIGS. 1a and 1b show schematically a longitudinal cross-section through the interior of previous aerofoil portions of a blade or vane. Each cross-section contains the leading edge L and trailing edge T of the aerofoil portion. In the FIG. 1a arrangement, air (indicated by arrows) is bled into the aerofoil section at an inlet in approximately the radial direction, travels along a radially extending passage, and exhausts from the trailing edge at about 90° to the radial direction. The peak thermal load is generally towards the centre of the aerofoil is, as indicated in FIG. 1a. FIG. 1b shows another arrangement in which the coolant passage forms a loop around a fence F, with the intention of directing a larger portion of the cooling air to the area of greatest thermal load.
In the prior art, the coolant inside the aerofoil flows axially and/or radially. Axial flow is predominantly in the same direction as the mainstream around the aerofoil (i.e. from leading-edge to trailing-edge). The coolant is then ejected into the mainstream through the trailing-edge or film cooling holes. As the coolant is fed to one of the near end-wall regions of the blade/vane, this results in the coolant with highest cooling potential (i.e. at the lowest temperature) being fed to the area of the blade/vane with the lower cooling requirement.
Further, the maximum wall temperature is a primary factor in determining the life of the aerofoil. Aerofoil cooling designs therefore attempt to minimise the maximum wall temperature. The maximum wall temperature typically occurs at midspan at the trailing-edge (where the external film is least effective) and the leading-edge (where mainstream stagnation temperatures are high).
There are many types of trailing-edge geometry in the prior art; for example, impingement systems that lead on to pedestal banks. These systems aim to offset elevated levels of external heat load at the trailing-edge of the aerofoil by a corresponding increase in internal heat pick-up. In most aerofoil designs, the wall temperature in the trailing-edge region increases to a maximum at the trailing-edge overhang (where the coolant has been ejected and is mixing into the mainstream). Because of internal heat pick up, the internal coolant increases in temperature through the trailing-edge system. The ability of the coolant to pick up heat diminishes as the coolant temperature increases. Therefore, the coolant is least effective where cooling is needed most: in the trailing-edge overhang region.
The present invention aims to at least partially overcome the limitations discussed above.